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8/10/2019 ACD 2510 Conceptual Aircraft Design Report
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MSRSAS - Postgraduate Engineering and Management Programme - PEMP
Aircraft Conceptual Design
i
FINAL REPORT
Module Code ACD 510
Module Name Aircraft Conceptual Design
Course M.Sc in Aircraft Design
Department Automotive and Aeronautical Engg.
Name of the Student ACD FT-11
Batch Full-Time / Part-Time 2011.
Module Leader Dr. H K Narahari
POSTGRADUATEENGIN
EERING
ANDMANAGEMENTPROGRA
MME(PEMP)
M.S.Ramaiah School of Advanced StudiesPostgraduate Engineering and Management Programmes(PEMP)
#470-P Peenya Industrial Area, 4th Phase, Peenya, Bengaluru-560 058
Tel; 080 4906 5555, website: www.msrsas.org
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Aircraft Conceptual Design
Declaration SheetStudent Name ACD FT-11
Reg. No
Course MSc [Egg] Aircraft Design Batch Full-Time 2011.Batch FT 11
Module Code ACD 510
Module Title Aircraft Conceptual Design
Module Date 06.08.2012 to 1.09.2012
Module Leader Dr. H K Narahari
Extension requests:Extensions can only be granted by the Head of the Department in consultation with the module leader.
Extensions granted by any other person will not be accepted and hence the assignment will incur a penalty.
Extensions MUST be requested by using the Extension Request Form, which is available with the ARO.A copy of the extension approval must be attached to the assignment submitted .
Penalty for late submissionUnless you have submitted proof of mitigating circumstances or have been granted an extension, the
penalties for a late submission of an assignment shall be as follows:
Up to one week late: Penalty of 5 marks
One-Two weeks late: Penalty of 10 marks
More than Two weeks late: Fail - 0% recorded (F)
All late assignments: must be submitted to Academic Records Office (ARO). It is your responsibility to
ensure that the receipt of a late assignment is recorded in the ARO. If an extension was agreed, the
authorization should be submitted to ARO during the submission of assignment.
To ensure assignment reports are written concisely, the length should be restricted to a limit
indicated in the assignment problem statement. Assignment reports greater than this length may
incur a penalty of one grade (5 marks). Each delegate is required to retain a copy of the
assignment report.
DeclarationThe assignment submitted herewith is a result of my own investigations and that I have conformed to the
guidelines against plagiarism as laid out in the PEMP Student Handbook. All sections of the text and
results, which have been obtained from other sources, are fully referenced. I understand that cheating and
plagiarism constitute a breach of University regulations and will be dealt with accordingly.
Signature of the student Date
Submission date stamp(by ARO)
Signature of the Module Leader and date Signature of Head of the Department and date
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Abstract____________________________________________________________________________
Aircraft design mainly depends on the existing historical data. Drastic changes in design cannot
be made hence it is usually the up gradation or the enhancement of the existing design. This
report is such an effort to Design of a Passenger Aircraft with a Range of 3000 nautical miles
with a passenger Capacity of 80. This report is made by a group of five members working on
different areas of design.
The first Chapter deals with the initial weight estimation which is usually the first step in the
design process of any aircraft. This is only a rough estimate rather than detailed weight
estimation because the weights must be varied in order to meet various customer requirements
in the later part of the design phase. This is because the design is a compromise of several
parameters.
Chapter 2 deals with the design of wing and high lifting devices. Various parameters like Wing
area, aspect ratio, airfoil, geometry, sweep angle, taper ratio is actually assumed based on the
historical data available. This is a good starting point because a lot of time is saved and most
importantly near close or meaningful results can be obtained in first set of assumption itself.
CFD analysis is carried out with the final geometry to validate the results obtained. Similar
procedure is carried out with flaps deflected at a particular angle calculated using Javafoil in
order to obtain max Cl needed as per the requirement.
Chapter 3 deals with the fuselage design which includes the entire layout and seating
arrangement which usually depends on the number of passengers. Several considerations are
carried out in design process to meet all FAA regulation like vision of the pilot and space
between the seats and also the dimension of the seats.
Chapter 4 deals with the selection of a propulsion system and the integration of the same. This
process depends upon the gross weight and the thrust requirements.Chapter 5 deals with the design of horizontal and vertical tail and various performance
parameters are calculated to check it meets with the customer requirement.
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Aircraft Conceptual Design
Contents____________________________________________________________________________
Declaration Sheet ......................................................................................................................... iiAbstract ....................................................................................................................................... iii
Contents ........................................................................................................................................iv
List of Figures ..............................................................................................................................vi
List of Nomenclature ....................................................................................................................ix
CHAPTER 1 .............................................................................................................................. 10
1.1 Introduction ....................................................................................................................... 10
1.2 Initial Weight Estimate ...................................................................................................... 10
1.2.1 Payload Weight .......................................................................................................... 11
1.2.2 Crew Weight ............................................................................................................... 11
1.2.3 Empty weight Fraction ............................................................................................... 11
1.2.4 Fuel weight Fraction ................................................................................................... 12
2.1 Wing Design ...................................................................................................................... 16
2.1.1 Wing loading .............................................................................................................. 17
2.1.2 Aspect Ratio ............................................................................................................... 17
2.1.3 Wing Sweep and Taper .............................................................................................. 18
2.1.4 Wing Geometry and Planform ................................................................................... 19
2.1.5 Number of wings ........................................................................................................ 20
2.1.6 Wing vertical location on the fuselage [1] ................................................................. 20
2.1.7 Steps for selection of the Airfoil for the wing [4] ...................................................... 23
2.1.8 Wing Twist ................................................................................................................. 27
2.2 Wing high lift devices ....................................................................................................... 30
2.2.1 Calculation of Takeoff and landing distance .............................................................. 31
2.2.2 Javafoil ....................................................................................................................... 32
2.2.3 Finite 3D Wing ........................................................................................................... 34
3.1Fuselage layout [6] ............................................................................................................. 36
3.2 Fuselage nose section ........................................................................................................ 38
3.3 Fuselage Mid-section ........................................................................................................ 39
3.4 Fuselage tail section .......................................................................................................... 39
3.5 Galley and Toilet configuration ........................................................................................ 40
3.6 Passenger loading and emergency exits ............................................................................ 40
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3.7 Structural consideration ..................................................................................................... 42
3.8 Weight estimation for fuselage .......................................................................................... 43
3.9 Landing gear layout: .......................................................................................................... 45
4.1Propulsion System .............................................................................................................. 474.2Engine Selection ................................................................................................................. 47
4.3 Why to go for CFM56-5B3 Engine ................................................................................... 49
4.4 Engine integration ............................................................................................................. 51
5.1 Empennage ........................................................................................................................ 52
5.2 Empennage types ............................................................................................................... 52
5.3 Empennage design ............................................................................................................. 53
5.4 Tail geometry .................................................................................................................... 53
5.5 Tail sizing .......................................................................................................................... 55
5.6 Tail layout .......................................................................................................................... 56
6.1 Empty weight build up and C.G location .......................................................................... 59
6.2 Performance ....................................................................................................................... 61
7.1 Assembly ........................................................................................................................... 64
8.1 Conclusion ......................................................................................................................... 65
REFERENCES .......................................................................................................................... 66
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List of Figures____________________________________________________________________________
Figure1. 1 Empty weight fraction Vs Wo[1] ............................................................................... 11
Figure1. 2 Mission Profile ........................................................................................................... 12
Figure1. 3 Aircraft type based on Range[2] ................................................................................ 13
Figure1. 4 Selection of SFC[1] .................................................................................................... 13
Figure1. 5 Selection of L/D Ratio[1] ........................................................................................... 13
Figure1. 6 Selection of L/D based on type of Aircraft[2] ........................................................... 14
Figure1. 7 No. of Passenger Vs MTOM[3] ................................................................................. 15
Figure1. 8 Range Vs MTOM/Passenger[3] ................................................................................. 16
Figure2. 1 Selection of Wing Loading[2][3] ............................................................................... 17
Figure2. 2 Selection of Aspect Ratio[2] ...................................................................................... 18
Figure2. 3 Mach no Vs Leading edge sweep angle[1] ................................................................ 19
Figure2. 4 Maximum and ideal lift co-efficient plot [5] ............................................................. 25
Figure2. 5 Cl Vs Cd plot for 631-412 [5] .................................................................................... 26
Figure2. 6 Airfoil created using ICEM CFD ............................................................................... 28
Figure2. 7 CATIA model of the wing (isometric view) .............................................................. 28
Figure2. 8 Top view of the wing ................................................................................................. 29
Figure2. 9 Domain around the wing ............................................................................................ 29
Figure2. 10 3-D domains around the wing .................................................................................. 29
Figure2. 11 Unstructured mesh for the wing ............................................................................... 30
Figure2. 12 Creation of Airfoil geometry ................................................................................... 32
Figure2. 13 Flap deflection ......................................................................................................... 33
Figure2. 14 Computation of Cl .................................................................................................... 33
Figure2. 15 Generation of curves using ICEM CFD from point data ......................................... 34Figure2. 16 CATIA model of 3D wing ....................................................................................... 34
Figure2. 17 Domain sketch ......................................................................................................... 35
Figure2. 18 Domain surface ........................................................................................................ 35
Figure2. 19 Mesh generated in ICEM CFD ................................................................................ 35
Figure3. 1 Seat pitch and height. [3] ........................................................................................... 36
Figure3. 2 Fuselage Interior detail (mm). .................................................................................... 37
Figure3. 3 Fuselage Plan view (mm). .......................................................................................... 37
Figure3. 4 Fuselage side view (mm). .......................................................................................... 38
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Figure3. 5 Nose section layout. [6] ............................................................................................. 38
Figure3. 6 Pilot Vision designed in Catia (mm). ......................................................................... 38
Figure3. 7 Pilot seat dimensions as per FAR. [3] ........................................................................ 39
Figure3. 8 Galley and toilet layout. [6] ....................................................................................... 40Figure3. 9 CATIA model for fuselage view-1 ............................................................................ 41
Figure3. 10 CATIA model fuselage view-2 ................................................................................ 41
Figure3. 11 Mesh geometry of fuselage and domain. ................................................................. 42
Figure3. 12 Mesh geometry of fuselage. ..................................................................................... 42
Figure3. 13 Stringers layout. [1] ................................................................................................. 43
Figure3. 14 Mass fractions for rapid mass estimation. [3] .......................................................... 44
Figure3. 15 Landing gear forward retract. [8] ............................................................................ 45
Figure3. 16 Nose landing gear load calculation. [8] ................................................................... 46
Figure3. 17 Tires used in typical aircraft. ................................................................................... 46
Figure 4. 1 Effect of flight speed on engine efficiency. [6] ........................................................ 48
Figure 4. 2 Effect of Mach number and specific thrust on thrust lapse rate. [6] ......................... 48
Figure 4. 3 Inlet locations-podded engines. [1] ........................................................................... 51
Figure 4. 4 Location of the nacelle compared to the wing .......................................................... 51
Figure 5. 1 Tail variations [1] ...................................................................................................... 52
Figure 5. 2 Tail aspect ratio and taper ratio for various types of aircraft [1] .............................. 54
Figure 5. 3 Tail volume coefficient [1] ....................................................................................... 55
Figure 5. 4 NACA 0012 airfoil coordinates with deflected control surface ............................... 57
Figure 5. 5 Velocity distribution and Cl values with deflected control surface .......................... 58
Figure 5. 6 NACA 0012 airfoil with deflected control surface in CATIA ................................. 58
Figure 5. 7 CATIA model-Horizontal tail ................................................................................... 59
Figure 5. 8 CATIA model-Vertical tail ....................................................................................... 59
Figure 6. 1 Approximate empty weight build up and c.g. location [1] ....................................... 60
Figure 7. 1 Front View ................................................................................................................ 64
Figure 7. 2 Top view ................................................................................................................... 64
Figure 7. 3 Side view ................................................................................................................... 64
Figure 7. 4 Isometric view ........................................................................................................... 65
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List of Tables
____________________________________________________________________________
Table1. 1Iterations ....................................................................................................................... 16
Table1. 2 2D Clincrement for leading edge flaps[2] .................................................................. 31
Table2. 1 Different set of airfoils selected from the graph [5] .................................................... 25
Table2. 2 Geometric twist for different aircrafts ......................................................................... 27
Table3. 1 Typical seat width and pitch for different class of travel. [6] ..................................... 36
Table3. 2 Typical guidelines for fuselage front and aft closure ratio. [3] ................................... 37
Table3. 3 Emergence exits requirements. [6] .............................................................................. 40
Table3. 4 Dimensions for types of exits. [6] ............................................................................... 41
Table3. 5 Mass amended for different configuration. [6] ........................................................... 43
Table 4. 1 Specifications and applications of CFM56-5 series engines. [10] ............................. 50
Table 5. 1 Design data of wing and fuselage .............................................................................. 54
Table 5. 2 Geometrical data supposed for tail design ................................................................. 54
Table 5. 3 Tail Design configuration ........................................................................................... 57
Table 6. 1 Empty weight build up and c.g. location .................................................................... 61
Table 6. 2 Derived data from design ........................................................................................... 62
Table 6. 3 Designed aircraft characteristics ................................................................................ 63
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List of Nomenclature____________________________________________________________________________
Acronym Expansion
C l Coefficient of lift
Cd Coefficient of Drag
TWR Thrust to Weight Ratio
WL Wing Loading
LDR Lift to Drag Ratio
Ct Specific Fuel Consumption
AR Aspect ratio
Wf
MTOM
Cdo
DSL
Fuel Weight
Maximum Takeoff weight
Zero lift Drag co efficient
Density at Sea Level
.
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CHAPTER 1
Design of a Passenger Aircraft with a Range of 3000 nautical miles with a passenger
Capacity of 80
1.1 Introduction
This report deals with the conceptual design of an aircraft. Any type of aircraft has the same
parts like fuselage, wing, engine etc but analyzing it a bit deeper it is evident that they are only
same by parts not on the geometry or the size. Fighter planes are smaller but much faster when
compared to that of the commercial transport plane. The root cause for such a big difference in the
design of aircraft is mainly due to the change in the functionality that has to be met. So in order to
start the conceptual design, proper understanding of the purpose of the aircraft to be designed is
required. Design in any field is a compromise of various factors, especially in the field of aircraft
where changing one influencing parameter affects many other influencing parameters. Since there
is a correlation between various influencing parameters the final modified design cannot be
achieved by changing only one influencing parameter. It is usually a combination of various
parameters to meet the final customer requirements and moreover other performance parameter as
per FAA (Federal Aviation Administration) regulation has to be satisfied for the aircraft to be
certified.An aircraft with a range of 3000 (nautical miles) and a passenger capacity of 80 comes under
the small Jet transport category. Looking into the available statistical data of similar type of aircraft
is regarded as the good starting point for a new design. The design process begins with the rough
estimate of the maximum take off mass. In this process several decisions has to be made on initial
assumptions to be used for calculations. Since most of the processes in aircraft design are iterative
in nature it is important to start with a meaningful assumption. By doing so reasonable results can
be achieved in a shorter span. This is done with the help of available statistical data.
1.2 Initial Weight Estimate
First step is to come up with a rough estimate of the total takeoff weight which is denoted by
Wo. It is the sum of crew weight, payload weight, fuel weight and empty weight.
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1.2.1 Payload Weight
Here the payload weight can be estimated with the number of passenger and the amount of baggage
that they are permitted to carry.
W payload = N * (Average weight of a passenger + Permitted baggage per individual)
Where N is the number of passengers and the initial assumption for average weight of an
individual passenger is taken as 80 kg (173.36 pounds) and the permitted baggage as 40 kg (88.14
pounds). So
W payload = 80 * (80+40) = 9600 kg (21165 pounds)
1.2.2 Crew Weight
Crew consists of a pilot, a co pilot, a flight engineer and for a passenger capacity of 80 we need two
air hostesses which all together make a crew as five members and assuming the crew members are
allowed to carry a baggage of 40 kg. Then
W crew = 5 * (120) = 600 kg (1323 pounds)
1.2.3 Empty weight Fraction
Empty weight is denoted as We, rather than calculating the empty weight as such, empty weight
fraction is calculated. Empty weight fraction is the ratio between the empty weight and the total
takeoff weight. This can be calculated from the statistical data available.
Empty weight fraction = (We/ Wo) and is given by We/ Wo = A* WoC* Kvs
Figure1. 1 Empty weight fraction Vs Wo[1]
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Since the aircraft to be designed comes under Jet transport appropriate values for A and C are
selected. Therefore
(We/Wo) = 1.02 Wo-0.06 *
1.04
(We/Wo) = 1.0608 Wo-0.06
1.2.4 Fuel weight Fraction
Fuel weight as such cannot be estimated from statistical data because amount of fuel to be carried
varies according to the mission profile. The mission profile is usually given as the customer
requirement. Statistical data are available to find out the fuel fraction at various stages of the
mission profile hence the total fuel fraction can be estimated.
Figure1. 2 Mission Profile
Where
1-2 Warm up and Takeoff
2-3 Climb
3-4 Cruise
4-5 Loiter and descent
5-6 Landing phase
Let W1, W2, W3, W4, W5and W6be the weights after respective phases of the mission profile.
1.2.4.1 Calculation
Amount of fuel utilized for warm up and takeoff phase is estimated as 0.97
(W1/W0) = 0.97
Amount of fuel utilized for climb phase with respect to the amount of fuel left over after thetakeoff phase is estimated as 0.985
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(W2/W1) = 0.985
Fuel utilized for Cruise depends upon the range to be achieved and can be calculated usingRange equation
Therefore Wi-1/ Wi in our case it is given by (W3/W2) is given by
Where R is the Range, C is the specific Fuel Consumption, V is the velocity, L/D is the lift to
drag ratio
Figure1. 3 Aircraft type based on Range[2]
Range for this mission is 3000 nautical miles (18228341.6 ft)
Specific fuel consumption for a high bypass turbo fan engine for cruise was found to be 0.5
1/hr (0.0001389 1/s)
Figure1. 4 Selection of SFC[1]
The cruise velocity was assumed to be 0.85 M (845.58 ft/sec).
In a similar fashion the L/D ratio for jet engine at cruise is given by L/D = 0.866 (L/D) max.
The selection of L/D and specific fuel consumption comes from the statistical data
available. (L/D) max was assumed to be 16. So L/D for cruise = 0.866*16 = 13.9
Figure1. 5 Selection of L/D Ratio[1]
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Figure1. 6 Selection of L/D based on type of Aircraft[2]
Substituting all the values we get
(W3/W2) = e-0.2154
= 0.8062
In case the intended airport is closed due to some unavoidable reason like worst climatic
condition, the aircraft has to loiter for extra duration or can have an increased range to reach
the nearest airport. According to FAA regulation an additional fuel for loitering atleast for
30 min has to be provided. In this case the additional time for endurance is taken as 45 min
which includes both loiter and descent. Fuel ratio calculation for endurance is as follows.
It can be obtained from endurance formula.
Therefore Wi-1/ Wi in our case it is given by (W4/W3) is given by
Where
E is the endurance in seconds. In this case it is 45 min (2700 sec)
But again the selection of specific fuel consumption and L/D ratio is based on statistical data
but it varies because it is a loitering phase. Specific fuel consumption for jet engines during
loiter phase is estimated as 0.4 1/ hr (0.0001111 1/s)
L/D for jet engines for loitering phase is the L/D max hence the value is 16 from statistical
data which says L/D selection can lie between (15 - 18.2).
Substituting all the values we get
(W4/W3) = e-0.01874
= 0.9814
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Fuel fraction used for landing is estimated as 0.995
(W5/W4) = 0.995
Therefore W5/W0= (W1/WO) * (W2/W1) * (W3/W2) * (W4/W3) * (W5/W4)
= 0.97 * 0.985 * 0.8062 * 0.9814 * 0.995
W5/W0= 0.7521
But the Total fuel fraction (Wf/ Wo) is given by the formula
(Wf/ W
o) = 1.06 * (1- W
5/W
0)
(Wf/ Wo) = 1.06 * (1- 0.7521) = 0.2626
We know that
Simplifying the equation, we get
77
It is an iterative process so the initial guess should be reasonable so that a lot of time can be saved.
Figure1. 7 No. of Passenger Vs MTOM[3]
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From the Figure 1.3 a initial estimate of 40000 kg approximately 80000 pounds is taken
Table1. 1Iterations
No ofteration nitial Gess ofWo Obtained ValeWo1 80000 113148.22 113150 107162.13 107000 1080754 108000 107922.45 107900 1079376 107933 107932.6
So the initial estimate of Maimm takeoff mass is 48957 kg (107933 ponds). The allatedale an be ross heked ith the range VS MTOM / passenger
Figure1. 8 Range Vs MTOM/Passenger[3]
So for a range of 3000 nm and a passenger capacity of 80 we have MTOM as 80*600 we get 48000
kg which is comparable to the estimated value.
2.1 Wing Design
The different parameters that needs to be found during the wing design is
Wing area
Aspect ratio
Taper ratio
Root chord
Tip chord
Mean aerodynamic chord
Span
Number of wings
Wing position on the fuselage
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Airfoil cross section
Twist angle
Sweep angle
Dihedral angle Wing incidence angle for Cruise
The step followed in the selection of wing area is discussed here in a sequential manner
2.1.1 Wing loading
Once the weight estimation is done, the next step is the proper selection of wing loading. The
selection of wing loading depends upon the mission requirement of an aircraft. Wing loading is the
ratio of total takeoff weight to that of the wing area. It is denoted by W/S and its units are Kg/m2or
lbs/ft2. It is found that for a short/ medium range aircraft the wing loading is about 110 lbs/ft
2
(537.06 kg/m2).
W/S = 537.06 Kg/m2(110 lbs/ft
2)
Figure2. 1 Selection of Wing Loading[2][3]
2.1.2 Aspect Ratio
Aspect ratio is defined as the ratio between the wing span to the wing mean aerodynamic. The
selection of aspect ratio in conceptual design phase is mainly on the historical data but what leads to
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the selection of particular aspect ratio is based on its effects on the various flight features such as
aircraft performance, stability, control, cost, and manufacturability.
The effects of aspect ratio are
As the aspect ratio increases, the aerodynamic features of the three-dimensional wing suchas CL, 0, CLmax, CDmin will get closer to the two-dimensional airfoil properties. This is
because of reduction of the influence of wing tip vortex, so it is desired have high AR.
As the AR increases, the maximum lift co-efficient for a particular angle of attack increases,
because the wing effective angle of attack increases, so it is desired to have a high aspect
ratio wing.
As the aspect ratio increases the weight of the wing increases, this needs more stiffer wing,
hence more stress on the root to hold the wing, hence it is desired to have a short wing to
reduce the weight of the wing.
From statistical data it is found that the aircraft that comes under the jet transport category has an
optimum Aspect ratio which lies between (7 - 9.5) hence a approximated value of aspect ratio 9 is
assumed. It is denoted by A.R and is a dimensionless quantity.
Aspect ratio = b2/S = 9
Figure2. 2 Selection of Aspect Ratio[2]
2.1.3 Wing Sweep and Taper
The next step in the design process is to come up with the wing geometry which includes the Root
chord and tip chord dimension, taper ratio, Sweep angle. The sweep back angle is provided to
reduce the effective Mach number at the leading edge. By doing so the loss of lift associated with
supersonic flow can be reduced. The sweep angle usually depends on the cruise Mach number.
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Figure2. 3 Mach no Vs Leading edge sweep angle[1]
From Figure 1.6 for a design cruise Mach no of 0.85 the leading edge sweep angle is chosen as 30
degrees. Taper ratio is defined as the ratio of the tip chord to that of the root chord. Tapered wing is
used in order to reduce the lift induced drag. A commercial transport aircraft with swept back wings
has a taper ratio that lie between 0.2 - 0.3. The maximum permitted taper ratio (0.3) is taken in this
design. So
Sweep Angle = 30 degrees
Taper Ratio = 0.3
2.1.4 Wing Geometry and Planform
Wing geometry includes calculation of Wing span, tip chord length, root chord length, Mean
aerodynamic chord length and span wise location of mean aerodynamic chord. All these parameters
can be calculated by the previously assumed parameters. The assumptions were Wing area = 90 m2,
Aspect ratio = 9, Wing sweep = 30 degrees, Taper ratio = 0.3.
Wing span b =
. =
9 9 0
Aspect Ratio =
=
Root chord length Cr = 4.864 m
For calculating Tip chord length, We have Taper Ratio = 0.3
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Tip chord length = 1.459 m
Mean Aerodynamic chord length
Span wise location of mean aerodynamic chord
2.1.5 Number of wings
In the olden days because of manufacturing limitations, more number of wings was used to
generate the required lift, now with modern manufacturing technologies and with the development
of new materials such as Aluminum and its alloys and composite have enabled the manufacturing
of single with long span wing.
With the modern technologies and the materials, now a days only single wing is used.
2.1.6 Wing vertical location on the fuselage [1]
The parameter that could be found during the initial stage of the design is the vertical location of
the wing with respect to fuselage centerline. This parameter will directly affect the other parameters
such as the tail location, landing gear design etc
Generally there are four types of configurations available, they are High wing
Mid wing
Low wing
Parasol wing
As seen from the figure that, the most of the cargo aircrafts have a high wing, where as fighter
aircrafts will have the mid wing and the long rage passenger aircrafts will have the low wing. The
advantages and disadvantages of the different configuration are
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High wing
The high wing has some advantages as well as some disadvantages for a particular mission, they are
Advantages
It eases the loading and unloading of cargo in to and out of the aircraft, and the trucks tounload and load can move easily around the aircraft.
The clearance from the ground for this configuration is more compared to low wing, which
facilitates the installation of engine on the wing.
It facilitates the aircraft to take-off and land from the sea in case of amphibian aircrafts, thus
preventing the spilling of water in to the engine during take-off, which may shut down the
engine.
The wing will produce more lift compared to low wing and mid-wing, because part of the
fuselage near the connection between two parts of the wing also contributes for the lift
produced by the wing.
Since the CL produced by the wing the wing is high the aircraft can fly at a lower stall speed
compared to high and the low-wing.
Disadvantages
The aircraft in this configuration will have more frontal area, which increases the drag of the
aircraft.
The ground effect will be lower compared to low wing, this will influence on landing and
take-off distances.
Landing gear is longer if connected to the wing. This makes the landing gear heavier and
requires more space inside the wing for retraction system. This will further make the wing
structure heavier.
The wing will produce more induced drag (Di), due to higher lift coefficient.
A high will be structurally 20% more heavier than the low wing.
Low Wing
In this section, advantages and disadvantages of a low wing configuration are discussed
Advantages
The aircraft take off performance will be better compared with a high wing configuration
due to ground effect.
The pilot will have a better view above the horizon, since the wing the wing is below the
pilot
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There will be option for the landing gear retracting system in the wing as well as fuselage.
Landing gear will be shorter; this makes the landing gear system lighter and requires less
space inside the wing for retracting system.
The aircraft is lighter compared to the high wing structure, aircraft frontal area in this caseis less. Since the frontal area is less, it has less induced drag.
Disadvantages
The wing generates less lift, compared with a high wing configuration since the wing has
two separate sections because of this the wing has less induced drag.
Because of the first reason the aircraft will have higher stall speed compared with a high
wing configuration due to a lower CLmax because of that the take-off run is longer.
The dihedral effect by the wing is less compared to the high wing, thus the aircraft is
laterally dynamically less stable.
The wing below the pilot will obstruct the view of the pilot below the horizon.
Mid Wing
In general, the features of the mid-wing configuration stand between features of high-wing
configuration and features of low-wing configuration. Some of the new features of a mid-wing
configuration are as follows:
The aircraft structure is heavier, due to the necessity of reinforcing wing root at the
intersection with the fuselage.
The mid wing is more expensive compared with high and low-wing configurations.
The mid wing is more attractive compared with two other configurations.
The mid wing is aerodynamically streamliner compared with two other configurations.
The strut is usually not used to reinforce the wing structure.
The pilot can get into the cockpit using the wing as a step in a small GA aircraft.
The mid-wing has less interference drag than low-wing and high-wing.
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2.1.7 Steps for selection of the Airfoil for the wing [4]
The design of airfoil is time consuming as well more cost is involved, for a conceptual design it is
better to select the best available airfoil from the data base. The steps involved in selection of the
airfoil are as follows.
The cruise altitude considered is 12 km at which the air speed is given by, temperature at 12 km is
216
= 7
The cruise number given for the design is 0.85 mach at an altitude of 12 km, therefore the cruise
velocity is given by
Vcruise = 249.8 m/s
First step is to determine the average weight in the cruising flight
7
Where Wiis the initial aircraft weight at the beginning of cruise and Wf is the final aircraft weight
at the end of cruise.
Calculation of aircraft ideal lift co-efficient (CLc).
In the cruising flight the aircraft weight is equal to the lift force.
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C= 2 VS
C=.
..
C= 2425299.810.3194249.8 90 =0.465Where V is the aircraft cruise speed, is the air density at cruising altitude, and S is the wing
planform area.
Calculate the wing cruise lift co-efficient (CLcw).
We consider that the wing the only component responsible for the generation of lift, but other
aircraft components such as tail, fuselage etcwill contribute to the total lift negatively or
positively. Thus the relation between aircraft cruise lift coefficient and wing cruise lift coefficient is
a function of aircraft configuration. The contribution of fuselage, tail and other components will
determine the wing contribution to the aircraft lift co-efficient. In the preliminary design phase
where the other components are been decided, then the following relation is used to calculate the
Wing cruise lift co-efficient.
C = .C = 0.4650.95 =0.489
Later in the design process when the other components are decided, this should be validated by
CFD simulations.
Calculation of wing airfoil ideal lift coefficient (Cli).
The wing is a 3-dimensional body whereas the airfoil is 2-D section, therefore the airfoil ideal lift
coefficient is different from the 3-D wing because the wing has a finite span and different chord
lengths and sweep angle results in this variation from the airfoil lift co-efficient, this variation can
be approximated using the relation.
C = C0.9
C = .. =0.54
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As per the statistical data the thickness to chord ratio generally used is 10-12%, but the cruise Cl
0.54 is not achieved in any of the airfoil with this thickness range with in the drag bucket. So the
wing area is modified to decrease the Cl cruise.
So the calculation are redone to adjust the Cl by considering wing area as 120 m2
, then the valueslift co-efficient values are
=
1Later in the design process when the other components are decided, this should be validated by
CFD simulations
Table2. 1 Different set of airfoils selected from the graph [5]
No NACA Cdmin Cmo s
(deg)
o
(deg)
Stall
quality
1 631-412 0.006 -0.08 14 -1.5 moderate
2 641-412 0.004 -0.040 12 -1.2 Moderate
3 651-412 0.004 -0.060 12 -1.5 Sharp
4 652-415 0.0035 -0.028 14 -1.2 Sharp
5 642-415 0.0045 -0.028 16 -1.3 Moderate
Figure2. 4 Maximum and ideal lift co-efficient plot [5]
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From the ideal lift co-efficient and the maximum lift co-efficient values required, the airfoil
with corresponding values of lift co-efficient are selected from the figure, where the figure
represents the collection of airfoils with different lift co-efficient values. If there is no airfoil ofparticular values then the airfoil that is nearest to the design point is selected.
If the wing is designed for a high subsonic passenger aircraft, select the thinnest airfoil (the
lowest (t/c max). The reason is to reduce the critical Mach number (Mcr) and drag-divergent9 Mach
number (Mdd). This allow the aircraft fly closer to Mach one before the drag rise is encountered. In
general, a thinner airfoil will have a higher Mcr than a thicker airfoil.
Figure2. 5Cl Vs Cd plot for 631-412 [5]
We can notice that the 631-412 is the airfoil which has the maximum Cl which is equal to the
calculated value and the cruise Cl of 0.41 according to the calculation, and the stall is moderate,
which is acceptable so 631-412 airfoil is chosen for the design.
Since the cruise Cl obtained is 0.41 to achieve that the angle of attack of 1.50 is required so the
wing is set at 1.50 angle for cruise. The result is validated using CFD
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2.1.8 Wing Twist
Twist is the difference in angle of attack between wing tip and the root, if the wing tip has lower
angle of attack than the root, the wing is said to have negative incidence and if the wing root is at
lower angle of attack then the tip, the wing is said to have positive incidence. In most of the cases
negative incidence is employed, the main reason for it is to have stall at the root first than at the tip,
because in case of stall if the root stalls first, then the pilot can still have a control on the ailerons to
control the aircraft, since the tip is at negative incidence, it will lower the overall lift produced by
the wing. There are two types of twists employed, they are
Aerodynamic twist
Geometric twist
If the different airfoil cross-section used in the root and tip, which will have different zero lift angle
of attack then, it s called as aerodynamic twist. If the tip and root have the same airfoil cross section
and if the incidence is not same then, it is referred to as Geometric twist.
In most of the cases aerodynamic twist is employed because it is easy to manufacture, where as
geometric twist is difficult to manufacture. The negative incidence of -1 to -4 is used in most of the
aircraft, because if the negative incidence is more then it decreases the overall lift produced by the
wing and the twist is also used to obtain the elliptical lift distribution on the wing.
In case of conceptual design phase, the twist is decided based on the historical data available and in
the later stage it can refined based on numerical calculations.
Table2. 2 Geometric twist for different aircrafts
No Aircraft MTOW
(lb)
Wing incidence at
root (iw) (deg)
Wing angle at
tip (deg)
Twist
(deg)
1 Fokker 50 20,800 +3.5 +1.5 -2
2 Cessna 310 4,600 +2.5 -0.5 -3
3 Cessna
Citation I
11,850 +2.5 -0.5 -3
4 Beech King
Air
11,800 +4.8 0 -4.8
5 Beech T-1A
JawHawk
16,100 +3 -3.3 -6.3
6 Beech T-34C 4,300 +4 +1 -3
7 Cessna
StationAir 6
3,600 +1.5 -1.5 -3
8 Gulfstream IV 73,000 +3.5 -2 -5.5
9 Northrop-
Grumman E-
2C Hawkeye
55,000 +4 +1 -3
10 Piper 11,200 +1.5 -1 -2.5
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Cheyenne
11 Beech Super
King
12,500 + 3o 48' -1o 7' 4.55'
12 Beech starship 14,900 +3 -5 -3.5
13 Cessna 208 8000 +2o 37' -3o 6' -5o 31'14 Beech 1900D 16,950 +3o 29' -1o 4' -4o 25'
15 Beechjet
400A
16,100 +3 -3o 30' -6o 30'
16 AVRO RJ100 101,500 +3o 6' 0 -3o 6'
17 Lockheed C-
130 Hercules
155,000 +3 0 -3
18 Pilatus PC-9 4,960 +1 -1 -2
19 Piper PA-28-
161 Warrior
2,440 +2 -1 -3
As seen from the table 2.2 for the design weight of 48900, the similar class of aircraft is Northrop-
Grumman E-2C Hawkeye, for which the twist angle of -30is used, since the design weight is close
to it, for the initial design phase the twist angle of -30is being used.
Figure2. 6 Airfoil created using ICEM CFD
Figure2. 7 CATIA model of the wing (isometric view)
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Figure2. 8 Top view of the wing
Figure2. 9 Domain around the wing
Figure2. 10 3-D domains around the wing
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Figure2. 11 Unstructured mesh for the wing
2.2 Wing high lift devices
The airfoil selected may give a design Cl required for cruise. But the same amount of lift may not
be sufficient during the takeoff and landing condition. Hence it is necessary to design high lifting
devices such as flaps and leading edge slats. The lift force is denoted by L and is given by
The need for high lift devices can be explained through the formula itself. The first thing to be
noticed is that the cruise velocity and the takeoff or landing velocity is not same. The takeoff
velocity is less when compared to that of the cruise velocity. So in order to achieve the desired lift
during takeoff and landing two approaches can be formulated, one is by increasing the C l and the
other way by increasing the surface area. Cl can be increased by increasing the camber of the
airfoil by deflecting the trailing edge flaps or the leading edge slats. There are several types of high
lift devices are available like plain flap, slotted flap, double and triple slotted flap, fowler flaps,
leading edge Krueger flap, slotted leading edge flap (slats) can be used. By using a plain flap, only
the camber of the airfoil can be changed but by using fowler flaps both the effective wing area and
also the camber can be varied. This is achieved by the extension or protrusion of the trailing edge
through some distance and is then deflected. Higher deflection can be achieved in case of fowler
flaps when compared to that of the split flaps because split flaps are prone to flow separation at
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higher deflection of the trailing edge but whereas in case of fowler flaps the flow separation does
not occur even at a higher angle of deflection because the effective chord length is also increased.
. Delta Clmax values for different type of leading edge flaps are
Table1. 2 2D Clincrement for leading edge flaps[2]
S.No Type Cl max
1 Fixed Slot 0.2
2 Leading Edge Flap 0.3
3 Kruger Flap 0.3
4 Slats 0.4
The design cruise Cl is 0.44 and if the leading edge slats are used the C l max value will be 0.8
which will also be not sufficient for takeoff and landing purpose. It is roughly estimated that the Cl
max required for takeoff and landing should be somewhere around 1.6 to 2.2 for long/ medium
range aircraft. Hence the desired Cl cannot be achieved just by using only one high lift device but
is usually a combination of both the leading edge slats and the trailing edge flaps. It is roughly
estimated that the leading edge flap deflection is usually 30 to 40 degrees. Since it is only a
conceptual design phase the normal split flap is only considered for initial computation rather than a
slotted flap. The most common flap chord length is 0.25 C from the trailing edge where C is theairfoil chord.
2.2.1 Calculation of Takeoff and landing distance
Landing and takeoff run is usually specified in terms of ground roll. A initial rough approximate of
takeoff distance is given by the formula [2]
Sg
=.(/)
Considering the design cruise Cl of 0.44 (i.e.) without using high lift devices.
Sg=
Sg = 12,726 ft (3,878 m)
So it is clearly evident that the takeoff distance is too large approximately 4 km. The airports cannot
afford to such a big runway. Moreover the takeoff and landing distance comes under customer
specifications and in most cases short take off and landing distance is preferred. The aircraft maynot be certified by FAA if the design does not meet the FAA regulations for takeoff and landing. As
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previously stated the Cl max required to meet takeoff and landing should be somewhere around 1.6
to 2.2 for long/ medium range aircraft. So assuming the required Cl max of 2.1, the landing distance
obtained will be 2666.42 fts (812.72 m) which is acceptable. So the takeoff distance with the
implementation of high lift devices is
Sg = 812.72 m
The rough estimate for landing distance for initial calculation is given by the formula [2]
SL= 118 (LP) + 400
Where LP is the landing parameter and is given by
LP =
Substituting the known values in the formula the landing distance is calculated as
SL = 29,900 ft (9113.52 m) for Cl of 0.44 without using high lift devices.
SL =6530 ft (2005.58 m) for Cl of 0.44 with the implementation of high lift devices.
These calculations clearly reveal the importance of lifting devices.
2.2.2 Javafoil
Figure2. 12 Creation of Airfoil geometry
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Javafoil is an applet which is available in public domain that can be used to come up with new
airfoil geometry as per the customer requirement. This applet is widely utilised not only to create
the required airfoil co-ordinates but it also provides options to vary the flap deflection angle andalso the location of percentage of flap with respect to the airfoil chord.
Figure2. 13 Flap deflection
The smooth option can be used to smoothen the curve when the flap is deflected.
Figure2. 14 Computation of Cl
It is found that the created airfoil generates a C l of 0.464 at zero degree angle of attack which is
bit higher than the design Cl of 0.44 but taking the 3 D effects into consideration the wing is
attached to the fuselage at an minimum angle not more than 2 degrees because the CL for wing
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will be less than 2D airfoil. It is also found that the desired flap deflection angle of 25.5 degrees is
required to achieve the Clmax value of (2.1). This is just an approximation of flap deflection
angle because the results may not be the same in case of a 3D wing hence a bit more deflection of
flap angle of 5 degree may be required to attain the Clmax
2.2.3 Finite 3D Wing
In order to study the variation between the 2D airfoil and a 3D wing, the obtained airfoil co
ordinates from Javafoil applet is converted into point data and is imported in ICEM CFD to
generate the airfoil curve and is then converted into IGES format and is taken to CATIA to create a
3 D wing geometry with all those taper ratio and sweep angle considered. Then the wing geometry
along with the domain is converted into IGES format and is imported to ICEM CFD. Meshing is
carried out in ICEM CFD and the mesh is read in Fluent. The resulting C l value will be less when
compared to that of the Javafoil case because of the 3D effects.
Figure2. 15 Generation of curves using ICEM CFD from point data
Figure2. 16 CATIA model of 3D wing
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Figure2. 17 Domain sketch
Figure2. 18 Domain surface
Figure2. 19 Mesh generated in ICEM CFD
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3.1Fuselage layout [6]
Structural requirements for pressurization mainly dictate the shape of the fuselage cross section.
Internal pressure loads are better handled by circular cross section by Hoop tension. This makes the
structure stronger and also lighter in comparison to other cross sectional shape. Non circular cross
section is more prone to bending stress for a shell structure. For a fuselage to maximize internal
volume usually interconnecting two or more circular section are considered.
Laying out the fuselage structure mainly depends on the payload specification. The number of
passenger dictates the width and length of the fuselage. The seat arrangement and number of aisle
helps in deciding the fuselage interior structures. The seat configuration considered in this design is
4 abreast single aisle and number of seats along the fuselage helps in fixing the length of the
fuselage. The length to diameter ratio for a given fuselage also needs to be considered at the time of
design since this helps in reducing drag. Low ratio increases the drag penalty where as high ratio
makes the fuselage long and thin which will adversely affect the structural stability of the airplane.
FAR rules have specified the minimum dimensions for different class of passenger seats and are
given in table 1.3.
Table3. 1 Typical seat width and pitch for different class of travel. [6]
Class Seat width (mm) Seat pitch (mm)
Charter 400-420 700-775
Economy 475-525 775-850
Business 575-625 900-950
First class 625-700 950-1050+
The seat width considered for the design is 500mm the seat pitch is 800mm and the aisle width as
500mm. the length of the fuselage for passenger seating is 800x20=16000mm. Fuselage consists of
a nose section, midsection barrel with constant cross section and aft-end closure.
Figure3. 1 Seat pitch and height. [3]
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Figure3. 2 Fuselage Interior detail (mm).
Table3. 2 Typical guidelines for fuselage front and aft closure ratio. [3]
Seating abreast Front fuselage
closure ratio. Fcf
Aft fuselage
closure ratio, Fca
Aft closure angle
(deg)
3 1.7-2 2.6-3.5 5-10
4-6 1.5-1.75 2.5-3.75 8-14
7 1.5 2.5-3.75 10-15
Figure3. 3 Fuselage Plan view (mm).
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Figure3. 4 Fuselage side view (mm).
Figure3. 5Nose section layout. [6]
3.2 Fuselage nose sectionThe fuselage nose section is from the tip of the nose end to the constant cross section of the mid
fuselage. This section holds the cockpit, flight deck, forward looking radar, nose undercarriage and
the windscreen. The fuselage length to diameter ratio considered is 1.5 as shown in figure 3.3. The
shape of the nose cone is designed such as to reduce drag.
Figure3. 6 Pilot Vision designed in Catia (mm).
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Figure3. 7 Pilot seat dimensions as per FAR. [3]
The pilot seat is standardized to have stress free condition for the pilot during takeoff and landing
with ample space to reduce fatigue to the pilot. The wind screen is designed to allow adequate
vision to the pilot for flight maneuver.
3.3 Fuselage Mid-section
This section holds the passenger seating, the length for this section depends on the number of
passenger to be accommodate. The seating are 4 abreast and 20 along with single aisle
configuration with a galley and toilet at both ends. Foldable seats one each at each end of the mid-
section for the crew is provided. The mid-section is 20148mm long as shown in figure3.3
3.4 Fuselage tail section
The tail section is shaped to provide smooth surface to reduce the drag. The tail section also
supports the tail surfaces and in some configuration the engine installation. The lower side of the
profile is tapered to 9deg to provide clearance for the aircraft during takeoff. The tail section length
to diameter ratio considered is 3. The overall fineness ratio of the fuselage is around ~10.
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3.5 Galley and Toilet configuration
Figure3. 8Galley and toilet layout. [6]
As per the FAR rules there should be one galley for 10-60 passengers and one toilet for 15-40
passenger. Since the design is for 80 passengers there are two galley and toilets provided at each
end of the midsection of the fuselage. The sizes considered for the galley is 762 mm x 914 mm and
for the toilet is 914mm x 914mm.
3.6 Passenger loading and emergency exits
FAR rules state that during emergency the plane needs to be evacuated within 90 second. This leads
to FAR guidelines for the number of emergency exits that is needed for different number of
passenger.
Table3. 3 Emergence exits requirements. [6]
Seats Emergency exit type
Less than Type I Type II Type III Type IV10 - - - 1
20 - - 1 -
40 - 1 1 -
80 1 - 1 -
110 1 - 2 -
140 2 - 1 -
180 2 - 2 -
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Table3. 4 Dimensions for types of exits. [6]
Type Dimension (mm)
Type I 610 x 1219
Type II 508 x 1118
Type III 508 x 914
Type IV 483 x 660
Type A (passenger or service loading door) 1067 x 1829
In consideration to the FAR rules two doors are provided at each end of the fuselage with Type-A
and two emergency exits one with Type I and one with Type III provided at each end of the
fuselage.
Figure3. 9 CATIA model for fuselage view-1
Figure3. 10 CATIA model fuselage view-2
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Figure3. 11 Mesh geometry of fuselage and domain.
Figure3. 12 Mesh geometry of fuselage.
3.7 Structural consideration
The fuselage structure should be able to take the bending moment, shear force, torsional loads and
the compressive loads due to self-weight, weight of wings and the weight of engine along with the
thrust force generated by the engine. The structure of the fuselage is constructed with a large
number of stringers distributed along the circumference of the fuselage which helps in resisting
bending of the fuselage. Bulkhead is provided at the ends of midsection of fuselage with frames
along the length of fuselage.
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3.8 Weight estimation for
Even though the weight of fuse
under carriage, initial guess c
considered at 7-12% of the m
formula [2] recommended for
fuselage mass. All the equati
correction factors need to be fa
Where LF= fuselage overall len
DF= fuselage diameter.
VD=aircraft maximum s
Table3
Configurat
For pressurized cabin
For fuselage mounted eng
For fuselage mounted mai
For large cargo door (etc.)
If free from structural disc
For the design considered the
m/s. the fuselage overall length
The fuselage body mass can be
_ = =1
Rapid mass estimation method
gineering and Management Programme - PEMP
Figure3. 13 Stringers layout. [1]
fuselage
age depends on the fuselage size, layout, an
an be assumed from historical data of si
ximum takeoff weight. For a better estima
civil aircraft (50-300 seats) to arrive at i
ns are for all metal (aluminum) constructi
tored in for different material or advanced li
th.
peed.
. 5 Mass amended for different configuration. [6]
on Mass to be a
Increase by
ines Increase by
n undercarriage Increase by
discontinuity Increase by
ontinuity Reduce by
light Mach # 0.85, hence the flight speed i
is 35898 mm, fuselage diameter is 3500 mm
estimated by the Howes formula as
^ ^889 kg.
an also be used to arrive at initial estimate f
43
location of engine and
ilar aircraft which is
ion one can use Howe
itial estimation of the
on only and necessary
hter material.
ended
8%.
4%.
7%.
10%.
4%.
at 340.3 x 0.85 = 289
.
r fuselage weight.
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Figure3. 14 Mass fractions for rapid mass estimation. [3]
Roskams [7] suggest few different ways of estimating the fuselage weight,
The general dynamic method,
=10.43(). 100 . 1000 . .Where Kinlet=1.25 for inlets in or on the fuselage, otherwise 1.0
qD=dive dynamic pressure in psf
L=fuselage length
D=fuselage depth
The Torenbeek method,
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=0.021 ( + ) . .
Here Kf=1.8 for a pressurized fuselage.
=1.07 for main undercarriage attached to the fuselage
= 1.1 for a cargo aircraft with rear door.
VD=design dive speed in knots equivalent air speed (KEAS)
LH_tail=tail arm of the H-tail
Sfus_gross_area= fuselage shell gross area.
In light of the fuselage construction it is very difficult to predict the weight since the weight
depends on the layout, type of material used and advancement of material technology for better
strength to weight ratio, which the initial estimate needs to be cross checked with detail calculation
and arrive at reasonably accurate estimation. Empirical formulas are available to get an accurate
estimate but are very time consuming since every structural component needs to be accounted for.
3.9 Landing gear layout:
The landing gear used is a tri cycle type. The layout of the landing gear is usually done at the end
after all the weight estimation is made for different section of the aircraft so that the CG is known
for the landing gear location. The landing gear location also depends on the tail-down angle
requirements suited for takeoff and landing attitudes, tipover and general airframe configurations.
The weight estimation also gives an idea whether to use two large wheels or 4 small wheels per
strut. For rough estimation 92% of gross weight is distributed on the main gear at aft CG condition
and 8% of the loads distributed on the nose landing gear at aft CG. The nose landing gear is placed
as far forward as possible to minimize the load on nose landing gear. The gear is designed to retract
forward to have a free fall capability. [8]
Figure3. 15 Landing gear forward retract. [8]
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Figure3. 16 Nose landing gear load calculation. [8]
The nose gear load is calculated as, [8]
Max static main gear load (per strut) = W (F-M)/2F
Max static nose gear load = W (F-L)/F
Min static nose gear load = W (F-N)/F
Where W is the gross weight,
For tire selection the nose gear dynamic load is necessary which is calculated as,
Max braking nose gear load = max static load + 10J. W/32.2F
Figure3. 17 Tires used in typical aircraft.
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From figure 3.17 the landing gear configuration used is a tri cycle type with two wheel per struts
with tire size 40 x 14 with 155 psi T type.
4.1Propulsion System
The turbofan is a type of air breathing jet engine that is widely used for aircraft propulsion. A
ducted fan which uses the mechanical energy from the gas turbine to accelerate air rearwards. The
ratio of the mass-flow of air bypassing the engine core compared to the mass-flow of air passing
through the core is referred to as the bypass ratio. Current high bypass ratio turbofan engines are
thermodynamically much more efficient than the early turbojet and low bypass types. This has been
largely brought about by the introduction of advanced technologies which have enabled turbine
blades to withstand high centrifugal loads whilst operating in gas temperatures considerably higher
than the melting point of the unprotected blade material. Most commercial aviation jet engines in
use today are of the high-bypass type, and most modern military fighter engines are low-bypass.
In passenger aircraft, efficiency is the main factor rather than performance, large aircraft the fuel
price accounts for some 30% of the aircraft direct operating costs. [3] By increasing the fuel
efficiency the amount of fuel carried will be lesser there by reducing the total weight. A more fuel
efficient engine will require less fuel to fly a given range and hence will lead to a lower take-off
weight.
We are designing 80 seater passenger aircraft having maximum takeoff mass is 48957 kg and range
of 3000 nm keeping this in mind we have to select the appropriate engine which suits the
requirements, the engine should weigh less weight so that maximum takeoff mass will not alter so
much and the engine should have better efficiency as the engine efficiency is one of the key factor
in passenger aircraft as efficiency decreases operating cost increases in order to decrease the
operating cost we need to choose the engine which is having better efficiency.
4.2Engine Selection
The weight estimation can be made by using the statically data of similar plane and plotting the
results of MTOW with number of passenger. Similar plots can also be made with range against
MTOW per passenger and using this number one can arrive at estimating the maximum take-off
weight (MTOW). Figure 1.1 and table 1.1 shows this relation and is used to arrive at first estimate
for the given design.
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Figure 4. 1 Effect of flight speed on engine efficiency. [6]
Above shows the difference between turboprop, turbofan and propfan engines, we can see from the
graph that Turboprop engines are having higher efficiency than Propfan and Turbofan engines but
Turboprop engine efficiency is high at lesser mach number (mach 0.5-0.6), but we need the cruise
speed of 0.85, for this cruise speed Turbofan engines are having higher efficiency than Propfan and
Turboprop, so it will be efficient if we install Turbofan engine to our design.
Figure 4. 2 Effect of Mach number and specific thrust on thrust lapse rate. [6]
Our aircraft design has total mass of 48957 kg and we know that thrust to weight ratio of passenger
aircraft is in between 0.3 to 0.4.
= 48957 9.81 =0.3
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Required thrust = 144080 N = 144.08kN
So we need to choose the engine which gives minimum thrust of 144.08kN. By consideringall above conditions one can go for Pratt & Whitney PW6000 series or CFM International
CFM56-5 series engine.
CFM International CFM56-5 series engine which gives the thrust of more than 100kN, in
our design the CFM56-5B3 engine can be used which gives thrust of 150kN, has bypass ratio of 5.4
and overall pressure ratio of 35.5, CFM56-5B3 is a dual rotor, axial flow turbofan engine, the
integrated fan and booster (low pressure turbine) is driven by a 4 stage low pressure turbine. A
single stage high pressure turbine drives the 9 stage high pressure compressor, the two rotors are
mechanically independent of each other. Air entering the engine is divided into a primary (inner)
airstream and a secondary (outer) airstream. After the primary airstream has been compressed by
the LPC and HPC, combustion of the fuel in the annular combustion chamber increases the HPC
discharge air velocity to drive the high and low pressure turbines. An accessory drive system off the
N2 rotor drives engine and airplane accessory components. [9]
4.3 Why to go for CFM56-5B3 Engine
CFM56-5B3 engines gives more thrust than the PW6000 series.
CFM56-5B3 engines has higher bypass ratio than PW6000 series engines thus there is
increase in the efficiency.
In CFM56-5B3 engines there is option of a double-annular combustor that reduces
emissions (particularly NOx).
A new fan in a longer fan case, and a new low-pressure compressor with a fourth stage.
It has higher efficiency than PW6000 series.
The CFM56-5B Tech Insertion configuration provides operators with up to 1 percent
improvement in fuel consumption over the life of the product compared to the base CFM56-
5B engine.
Low carbon emission than PW6000 series engines.
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Table 4. 1 Specifications and applications of CFM56-5 series engines. [10]
( )
.
() 133.46 137.91 146.81 120.11 97.87 104.54 120.11 96.09 103.65
(/) 4194.65 4252.47 4350.85 3990 3638.6 3754.3 3990 3607.5 3741
5.5 5.5 5.4 5.7 6 5.9 5.7 6 5.9
(10.668 0.80 )
() 28.56 28.56 28.56 25.05 25.05 25.05 28.56 25.05 25.05
35.4 35.4 35.5 32.6 32.6 32.6 35.5 32.6 32.6
.
.
() 25.98 25.98 25.98 22.33 22.33 22.33 25.98 22.33 22.33
() 2.6 2.6 2.6 2.6 2.6 2.6 2.6 2.6 2.6
() 1.734 1.734 1.734 1.734 1.734 1.734 1.734 1.734 1.734
() 23.36 23.36 23.36 23.36 23.36 23.36 23.36 23.36 23.36
: 321 321 321 320 319 319 319 319 319
319
Specific Fuel Consumption of the engine is 1.00034*10-4
N/N-S. [11]
Lapse rate = 5.650 C/km
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4.4 Engine integration
The engine integration has a significant impact on the aircraft, affecting safety, structural weight,
flutter, drag, control, maximum lift, propulsive efficiency, maintainability, and aircraft growth
potential. Engines can be mounted on different wing positions,
Figure 4. 3 Inlet locations-podded engines. [1]
Engine location is influenced by many considerations including the interference between the
nacelle and the wing which increases drag. Consequently, nacelles must be sufficiently forward and
low to avoid drag increases. However, to minimize the weight of the landing gear and engine pylon,
a general rule is drawn, the nacelles are usually located as close to the wing lower surface as
possible, without causing undue heating of the wing by the engine exhaust.
Figure 4. 4 Location of the nacelle compared to the wing
From the literature survey, the engine can be placed below the wing as shown in the figure.
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From the data of thrust to weight ratio, the minimum required thrust is found to be 144.08kN for
our design and CFM56-5B3 engine is selected for the aircraft and clarification for choosing this
engine is discussed.
5.1 Empennage
The function of an empennage is to stabilize the aircraft and provide control moments
needed for maneuver and trim. The empennage consists of a horizontal and a vertical tail. Together
they stabilize an aircrafts pitch and yaw moments. Trim for a horizontal tail refers to the balancing
of moment created by the wing. For a vertical tail, trim force generated by it is largely unexploited,
since most aircrafts are axis symmetrical. But in the case of an engine failure the vertical tail must
provide for enough trim to sustain the aircraft stable. Though it is possible to build a tailless
aircraft, it often comes with greater compromises in weight, wing area, airfoil selection and narrow
centre of gravity range. The other major function of the tail is to provide control. The tail must be
sized so as to provide adequate control at all critical conditions. For a horizontal tail, this includes
control during takeoff and landing, low speed flight and transonic maneuvering. For a vertical tail,
engine out flight, spin recovery and maximum roll rate are vital control conditions.
5.2 Empennage types
Different types of tail variations are available for various aircrafts depending on their functionality.
Some variations in tail design are as shown in figure 5.1.
Figure 5. 1Tail variations [1]
Each type of tail has its uniqueness owing to its functionality. Some of the designs are explained as
follows.
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Conventional tail
About 70% of the aircrafts use a conventional tail. This is due to the fact that a conventional
tail will be able to produce the necessary stability and control at the lightest weight.
T tailThis design is very popular due to its high aerodynamic efficiency. Since the horizontal tail
is above, it avoids wing wake and propwash and also reduces fatigue in both structure and pilot. A
smaller area of the vertical and horizontal tails would suffice compared to a conventional tail to
produce the necessary moments. But it comes with a high weight penalty though.
Cruciform tail
The cruciform tail is a compromise between conventional tail and the T tail. It can function
even at high angles of attack like the T tail but with comparatively lesser weight penalty. However
no reduction in tail area can be made as in the case of a T tail.
H tail
In this design the vertical tail is positioned as such so as to have undisturbed flow of air at
high angles of attack. Also, in the case of twin engines the H tail is positioned as such to be in line
of propwash so as to have better control in the case of engine out.
V tail
The V tail is intended to reduce wetted area. With a V tail the horizontal and vertical tail
forces are the result of horizontal and vertical projection of forces exerted on the V surface. The
resulting wetted area of a V tail would be lesser than that of separate horizontal and vertical tails. V
tails offer reduced interference drag, but with some penalty in control surface actuation complexity
as the rudder and elevator controls must be blended to provide the proper movement of V tail
ruddervators. This also results in adverse roll-yaw coupling.
Hence for the purpose of commercial transport aircraft a conventional tail would be appropriate.
5.3 Empennage design
The tail design is quite similar to the design of a wing. However a smaller area of tail is
enough to compensate for the moments about the aircraft aerodynamic centre due to the distance
between tail and wing.
5.4 Tail geometry
The surface area required by the tail is directly proportional to the area of the wing. Hence
area of the tail cannot be determined without the area of the wing. However other geometric
parameters like aspect ratio and taper ratio are similar over a wide range of aircraft types. These are
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obtained through statistical data and can be adhered to at the initial stage of tail design. Statistically
obtained data of tail aspect ratio and taper ratio for various types of aircraft is as shown in figure
5.2.
Figure 5. 2 Tail aspect ratio and taper ratio for various types of aircraft [1]
Other geometric parameters are selected based on the following guidelines:
Tail thickness ratio is similar to wing thickness ratio as per historical guidelines
provided in the wing geometry section. Since the aircraft designed is a commercial
transport aircraft, the airfoil selected for the tail is NACA0012 which is similar to the
thickness of the wing airfoil.
Horizontal tail leading edge sweep is set to about 5 deg more than the wing sweep. It
enables the tail to stall after the wing and provides the tail with higher critical Mach
number. Hence loss of elevator effectiveness due to shock formation is avoided.
Vertical tail sweep varies from 35 to 55 deg for most aircrafts. This is to ensure that the
vertical tail has higher critical Mach number than the wing.
Data obtained from wing design and fuselage design of the given aircraft is as shown in table5.1
Table 5. 1 Design data of wing and fuselage
Wing (a